Aeronautics and astronautics – Aircraft – heavier-than-air – Airplane and helicopter sustained
Reexamination Certificate
2001-11-08
2002-12-24
Swiatek, Robert P. (Department: 3643)
Aeronautics and astronautics
Aircraft, heavier-than-air
Airplane and helicopter sustained
C244S017110, C244S039000, C416S22300B
Reexamination Certificate
active
06497385
ABSTRACT:
BACKGROUND OF THE INVENTION
1. Field of the Invention
The present invention relates to a rotor blade for a rotary-wing aircraft, and more particularly, to such a blade with an increased hover figure of merit as compared to known blades.
2. Description of Related Art
Hover efficiency is a crucial determining factor in rotor blade performance, strongly influencing both range and payload of a rotary wing aircraft such as a helicopter or tiltrotor aircraft.
Although rotor performance in hover is greatly influenced by the very complex and unsteady vortical wake generated beneath the rotor, much of the current methodology for rotor blade aerodynamic design is based on momentum and blade element theories that do not account for the effect of this vortical wake. Classical vortex theory has also been applied to this problem, but by necessity it requires the use of simplifying assumptions such as prescribing the position of the vortex wake, which limits the extent to which it improves over momentum theory for design applications.
Classical methods are reasonably accurate in predicting rotor hover performance. However, even a very small increase in hover efficiency can provide important improvements in rotary wing aircraft payload and range. Classical methods are not capable of calculating performance to the required level of accuracy to predict these small increases in hover efficiency.
A typical rotary wing helicopter H is depicted in FIG.
1
. It includes a fuselage FH to which is mounted landing gear LH, a main rotor RH and a tail rotor TH. The helicopter H has one or more powerplants PH to provide motive force to the main rotor RH and the tail rotor TH. The main rotor RH includes a plurality of rotor blades BH, with which the present invention is concerned, mounted to a rotor hub HH. It will be appreciated by those skilled in the art that the helicopter can have a different number of blades depending on the performance parameters it must satisfy.
U.S. Pat. No. 3,882,105, U.S. Pat. No. 5,035,577 and U.S. Pat. No. 6,000,911 are examples of prior art attempts to provide optimized rotor blades for such a helicopter. These patents focus on numerous geometrical properties of the blade, including twist. Blade geometric twist &thgr; determines a blade's geometric angle of attack at each spanwise location relative to the airflow approaching the blade, and thus the lift generated by the blade as it rotates. (By convention, positive values of &thgr; signify upward twist of the blade leading edge.) These patents state that increasing overall blade twist beyond that of previous configurations reduces the geometric angle of attack and the local lift near the blade tip in hover. This is said to diminish the strength of the vortex trailed from the tip of the blade as it rotates and thereby reduce vortex interference and induced power losses and increase aerodynamic efficiency.
Those skilled in the art understand that a given blade's figure of merit F
M
is a generally accepted indication of a rotor's hover efficiency. Figure of merit is defined as the ratio of minimum possible power required to hover, to the actual power required to hover. Thus, figure of merit compares the actual rotor performance with the performance of an ideal rotor. Johnson, W.,
Helicopter Theory
, Princeton Univ. Press (1980), pages 34-35 (hereinafter “Johnson”).
Mathematically, figure of merit can be expressed as follows:
F
M
=
0.7071
C
T
1.5
C
Q
where
Rotor
torque
coefficient
C
Q
=
Q
ρ
π
R
3
(
Ω
R
)
2
Rotor
thrust
coefficient
C
T
=
T
ρ
π
R
2
(
Ω
R
)
2
In the above equations:
Q=torque in pounds-feet,
T=thrust in pounds,
R=rotor radius in feet measured from the axis of rotation,
&OHgr;=rotor angular velocity in radians per second, and
&rgr;=density of air in slugs per cubic feet.
It will be appreciated that F
M
for a particular rotor is an indication of the ratio of the induced power required to produce a given amount of thrust if the air were uniformly accelerated through the rotor disc around the azimuth, to the actual total power (induced plus profile) required to produce the same amount of thrust with the actual rotor. Induced power is a consequence of the fact that the lifting force induced by the rotating blades is not directly vertical, and therefore has a component producing what is known to those skilled in the art as induced drag. Profile power is a consequence of the profile drag on the rotating blades.
U.S. Pat. No. 3,882,105 proposes changing the geometric twist of a conventional rotor blade near the tip. Geometric twist is the angle of a blade chord relative to a reference plane. The patent superimposes a span-wise twist distribution on an otherwise conventional rotor blade in a tip region of the blade (in the patent, outward to the blade tip from 71% to 88% of the blade span, depending on the number of blades in the rotor). The nomimal twist distribution disclosed in the patent provides a tip region with an ever-decreasing amount of incremental twist relative to the twist of the conventional blade, as shown in
FIG. 2
of the patent.
U.S. Pat. No. 3,882,105 also suggests that the blade at an even smaller region closer to the tip (outward to the tip from 80% to 96% of the blade span, again depending on the number of blades in the rotor) can be twisted in the opposite direction to increase blade twist. The patent specifically states that this relatively upward twist close to the tip is no greater than a maximum of 3.5°, and should never exceed an amount that maximizes lift-to-drag ratio in this region.
The only helicopter known actually to use a blade with upward twist in a tip region is Sikorsky's UH-60. The UH-60 blade has a tip region in which the twist angle decreases incrementally beginning at about 85% of the blade span. The region from about 93% to the tip has an increasing twist angle (that is, the blade twists back upward) about 2° or so. The UH-60 blade is discussed in Johnson, W. R., “Wake Model for Helicopter Rotors in High Speed Flight,” NASA CR 177507, November 1988, p. 260. It appears that the UH-60 blade's twist distribution is intended to follow the teachings of U.S. Pat. No. 3,882,105.
There is also prior art that suggests that sweeping the blade tip backward and drooping the blade tip downward (providing an anhedral angle) will further improve hover efficiency.
Classic momentum theory suggests that in hover mode the optimum twist angle distribution along the blade span is proportional to 1/r, where r is the location along the blade measured from the axis of rotation. However, using such a twist distribution in the inboard regions of the blade (that is, from approximately r/R=0.20 to r/R=0.75), increases vibration in forward flight. Accordingly, twist in that region of helicopter rotor blades is usually reduced to below that which is called for by classical momentum theory, resulting in a concomitant performance penalty.
In summary, even though helicopter blade performance has been improved over the years by approaches such as that disclosed in the above-mentioned patents and other prior art, rotor blade designers continue to seek further performance enhancements for the hover, forward flight and landing regimes of helicopters.
Those efforts are hampered by the complexity of the flow created by helicopter rotors and the difficulty in analyzing those flows. But as difficult as it is to optimize a particular helicopter rotor blade configuration, tiltrotor aircraft provide an even greater challenge to the blade designer.
This type of aircraft has rotary blades that enable it to take-off as a helicopter, and are then tilted to provide forward propulsive thrust in a mode of flight in which the aircraft operates as a propeller-driven airplane. An example of such an aircra
Quackenbush Todd R.
Wachspress Daniel A.
Continuum Dynamics, Inc.
David M. Quinlan P.C.
Swiatek Robert P.
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